In order to increase the efficiency and the performance of gas turbine engines so as to provide increased thrust-to-weight ratios, lower emissions and improved specific fuel consumption, engine turbines are tasked to operate at higher temperatures. As the higher temperatures reach and surpass the limits of the material comprising the components in the hot section of the engine and in particular the turbine section of the engine, new materials must be developed.
As the engine operating temperatures have increased, new methods of cooling the high temperature alloys comprising the combustors and the turbine airfoils have been developed. For example, ceramic thermal barrier coatings (TBCs) were applied to the surfaces of components in the stream of the hot effluent gases of combustion to reduce the heat transfer rate and to provide thermal protection to the underlying metal and allow the component to withstand higher temperatures. These improvements helped to reduce the peak temperatures and thermal gradients. Cooling holes were also introduced to provide film cooling to improve thermal capability or protection. Simultaneously, ceramic matrix composites were developed as substitutes for the high temperature alloys. The ceramic matrix composites (CMCs) in many cases provided an improved temperature and density advantage over the metals, making them the material of choice when higher operating temperatures were desired.
A number of techniques have been used in the past to manufacture turbine engine components, such as turbine blades using ceramic matrix composites. However, such techniques have resulted in difficulties related to the small features of gas turbine engine components for helicopter engines.
A number of techniques have been used in the past to manufacture turbine engine components, such as turbine blades using ceramic matrix composites. One method of manufacturing CMC components, set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350; incorporated herein by reference in their entirety and assigned to the assignee of the present invention, relates to the production of silicon carbide matrix composites containing fibrous material that is infiltrated with molten silicon, herein referred to as the Silcomp process. The fibers generally have diameters of about 140 micrometers or greater, which prevents intricate, complex shapes having features on the order of about 0.030 inches, such as turbine blade components for helicopter gas turbine engines, to be manufactured by the Silcomp process.
Other techniques, such as the prepreg melt infiltration (MI) process have also been used, however, the smallest cured thicknesses with sufficient structural integrity for such components have been in the range of about 0.03 inch to about 0.036 inch, since they are manufactured with standard prepreg plies, which normally have an uncured thickness in the range of about 0.009 inch to about 0.011 inch. With standard matrix composition percentages in the final manufactured component, the use of such uncured thicknesses results in final cured thicknesses in the range of about 0.03 inch to about 0.036 inch.
What is needed is a method of manufacturing CMC helicopter turbine engine components that permits the manufacture of features having a thickness in the range of about 0.015 inch to about 0.021 inch. In addition, a method of manufacturing CMC helicopter turbine engine components having features with a thickness less than about 0.021 inch is also needed.